The invention relates to the field of gyro referenced satellite systems. More particularly, the present invention relates to systems and methods for emulating gyros for improved reliability and lifetime of flight systems, as particularly described in U.S. Pat. No. 6,020,956 entitled Pseudo Gyro, Issued Feb. 1, 2000, and U.S. Pat. No. 6,263,264 entitled Pseudo Gyro with Unmodeled Disturbance Torque Estimations, both of which are here incorporated by reference as there fully set forth.
A hardware gyro is used to provide vehicular bus rotational rates, that is, the angular velocity data used for vehicular attitude reference and attitude control. In order to accomplish controlled attitude orientation and controlled angular velocity of the satellite, the control system requires accurate attitude data from the reference system and accurate angular velocity rate data from the gyro. The gyro operation is subject to errors and biases that are corrected by a low frequency filter such as a Kalman filter in the attitude reference system. The low frequency filter is any filter that receives measurement data and produces update data with a low frequency bandwidth within a high bandwidth of the system operation. Hence, the gyro is an integral and essential part of an attitude reference system and is necessary for the proper functioning of the attitude control system. Modern applications require accurate, high bandwidth gyroscopes with long lifetimes, high reliability, and low power usage.
The appendage measurement data, reaction wheel tachometer data, and external torque rod current data define changes in momentum of the bus. These appendage measurement data, reaction wheel tachometer data, and external torque rod current data are momentum data related to the momentum of the spacecraft and processed by the attitude control system. This momentum data is available for momentum computational purposes. Gyros have accordingly long been used successfully in satellite attitude reference systems to provide the vehicular angular velocity rates of the satellite for satellite attitude referencing and positioning control. However, gyros are expensive and inherently have limited life times with low reliability and dynamic error characteristics, leading to premature failure of satellite systems.
U.S. Pat. Nos. 6,020,956 and 6,263,264 teach the use of a pseudo gyro realized by a software process to emulate the functions of a hardware gyro. The pseudo gyro uses the principle of conservation of momentum to compute the bus rate by accounting for the momentum from external torques and the transfer between a satellite bus and the appendages and momentum storage devices such as reaction wheels. The external torques, the appendage measurements, and reaction wheel tachometer data are used by the pseudo gyro to calculate the bus angular velocity rate. Accurate relative position and rate information are typically available from the attitude and appendage controllers on-board the satellite to facilitate these computations. The pseudo gyro is used as an integral part of an attitude reference system of a space satellite, but can be applied to other flight and space systems by replacing the hardware gyros. This pseudo gyro operates as part of the control and reference systems now having higher reliability, longer life times, lower power consumption, and more accurate angular velocity rates within high bandwidth operations. The pseudo gyro may operate concurrently with on board gyros that are subject to failure limiting the life of a spacecraft. The pseudo gyro does not detect errors in the conncurrently operating hardware gyros.
Hardware fault detection and resolution is extremely important to satellite operators. Time wasted resolving hardware anomalies results in payload outages and often endangers the safety of the spacecraft. For this reason, many satellites now include automated fault detection and redundancy management software. Gyros are key satellite attitude determination and control sensors. Unfortunately, existing fault detection software can not reliably detect a gyro failure when less than four gyros are operating simultaneously and fault detection software can not isolate the faulty gyro when less than five gyros are operating simultaneously. When the gyros are not properly oriented, up to six gyros may be required. Typical satellite attitude control and determination requires only three properly oriented gyros. Operating more than three gyros is undesirable for power and lifetime considerations. Therefore, sophisticated fault detection software, requiring more than three gyros, is not generally applied to detect gyro faults. A gyro failure on a commercial communications satellite lead to the loss of nominal attitude control resulting in major television network outage for the United States. Several hours are required to identify that a gyro failure caused the anomaly, determine which of the three gyros was in error, and swap the redundant gyro for the failed gyro. These and other disadvantages are solved or reduced using the invention.
An object of the invention is to provide a reference system in a flight or space system using an emulated gyro receiving the momentum data for estimating of external torque and to provide an estimate of an error in the external torque for emulating a gyro.
Another object of the invention is to provide fault detection of hardware gyros.
A further object of the invention is to provide fault detection of momentum sensors.
The invention is directed to a pseudo gyro software process to detect failure of operating hardware gyros. The pseudo gyro is preferably used as an integral part of an attitude reference system of a space satellite, but can be applied to other flight and space systems by replacing the hardware gyros. This pseudo gyro operates as part of the control and reference systems now having higher reliability, longer life times, lower power consumption, more accurate angular velocity rates within high bandwidth operations. The pseudo gyro is used to monitor gyro performance for fault detection. The pseudo gyro is extended to provide a fault detection of hardware gyros. The improved pseudo gyro provides an on-orbit functionality check of hardware gyros, provides an on-orbit performance check of hardware gyros, identifies the specific gyros with anomalous performance, functions with any number of hardware gyros, provides long term performance trending of hardware gyros, identifies anomalous performance of hardware used in the pseudo gyro solution such as stepper motors, encoders, angular orientation sensors, tachometers, and reaction wheels. These and other advantages will become more apparent from the following detailed description of the preferred embodiment.